This application relates to a method of providing heat to a rotor bore in a gas turbine engine at certain times during operation of an aircraft.
Gas turbine engines are known and when used on aircraft typically include a fan delivering air into a bypass duct and into a compressor section. Air from the compressor is passed downstream into a combustion section where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors driving them to rotate.
Turbine rotors drive compressor and fan rotors. Historically, the fan rotor was driven at the same speed as a turbine rotor. More recently, it has been proposed to include a gear reduction between the fan rotor and a fan drive turbine. With this change, the diameter of the fan has increased dramatically and a bypass ratio or volume of air delivered into the bypass duct compared to a volume delivered into the compressor has increased. With this increase in bypass ratio, it becomes more important to efficiently utilize the air that is delivered into the compressor.
One factor that increases the efficiency of the use of this air is to have a higher pressure at the exit of a high pressure compressor. This high pressure results in a high temperature increase. The temperature at the exit of the high pressure compressor is known as T3 in the art.
There is a stress challenge to an increasing T3 on a steady state basis due largely to material property limits called “allowable stress” at a given maximum T3 level. At the maximum, a further increase in a design T3 presents challenges to achieve a goal disk life. In particular, as the design T3 is elevated, a transient stress in the disk increases because the radially outer portions of a high pressure compressor rotor (i.e., the blades and outermost surfaces of the disk or blisk), which are in the path of air, see the increased heat rapidly when T3 shoots up rapidly during a rapid power increase such as when the pilot increases power during a take-off roll. However, a rotor disk bore does not see the increased heat as immediately. Thus, there are severe stresses due to the thermal gradient between the disk bore and the outer rim region.
This thermal gradient challenge is greatest during the take-off of an aircraft engine and it is possible that the thermal stress in the disk is much greater than the stress due to the centrifugal force on the disk—particularly in the compressor where the blades are light. The engine has typically been at low speed or idle as the aircraft waits on the ground and then, just before take-off, the speed of the engine is increased dramatically. The thermal gradient stresses have led to the high pressure compressor often being operated at a lower pressure (and hence T3) than would be optimum.